U.S. patent application number 15/618289 was filed with the patent office on 2018-12-13 for multiple chamber rotating detonation combustor.
The applicant listed for this patent is General Electric Company. Invention is credited to Clayton Stuart Cooper, Arthur Wesley Johnson, Sibtosh Pal, Steven Clayton Vise, Joseph Zelina.
Application Number | 20180355822 15/618289 |
Document ID | / |
Family ID | 64563260 |
Filed Date | 2018-12-13 |
United States Patent
Application |
20180355822 |
Kind Code |
A1 |
Vise; Steven Clayton ; et
al. |
December 13, 2018 |
MULTIPLE CHAMBER ROTATING DETONATION COMBUSTOR
Abstract
The present disclosure is directed to a rotating detonation
combustion system for a propulsion system including a plurality of
combustors in adjacent arrangement along the circumferential
direction. Each combustor defines a combustor centerline extended
through each combustor, and each combustor comprises an outer wall
defining a combustion chamber and a combustion inlet. Each
combustion chamber is defined by an annular gap and a combustion
chamber length together defining a volume of each combustion
chamber. Each combustor defines a plurality of nozzle assemblies
each disposed at the combustion inlet in adjacent arrangement
around each combustor centerline. Each nozzle assembly defines a
nozzle wall extended along a lengthwise direction, a nozzle inlet,
a nozzle outlet, and a throat therebetween, and each nozzle
assembly defines a converging-diverging nozzle. A first array of
combustors defines a first volume and a second array of combustors
defines a second volume different from the first volume.
Inventors: |
Vise; Steven Clayton;
(Loveland, OH) ; Zelina; Joseph; (Waynesville,
OH) ; Johnson; Arthur Wesley; (Cincinnati, OH)
; Cooper; Clayton Stuart; (Loveland, OH) ; Pal;
Sibtosh; (Mason, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
64563260 |
Appl. No.: |
15/618289 |
Filed: |
June 9, 2017 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C 5/11 20130101; F02C
5/12 20130101; F05D 2220/80 20130101; F02K 7/06 20130101; F02K
7/067 20130101 |
International
Class: |
F02K 7/00 20060101
F02K007/00 |
Claims
1. A rotating detonation combustion system for a propulsion system,
the rotating detonation combustion system defining a radial
direction, a circumferential direction, and a longitudinal
centerline in common with the propulsion system extended along a
longitudinal direction, the rotating detonation combustion system
comprising: a plurality of combustors in adjacent arrangement along
the circumferential direction, wherein each combustor defines a
combustor centerline extended through each combustor, and wherein
each combustor comprises an outer wall defining a combustion
chamber and a combustion inlet, wherein each combustion chamber is
defined by an annular gap and a combustion chamber length together
defining a volume of each combustion chamber, and further wherein
each combustor defines a plurality of nozzle assemblies each
disposed at the combustion inlet in adjacent arrangement around
each combustor centerline, and wherein each nozzle assembly defines
a nozzle wall extended along a lengthwise direction, a nozzle
inlet, a nozzle outlet, and a throat therebetween, wherein each
nozzle assembly defines a converging-diverging nozzle, and wherein
a first array of combustors defines a first volume and a second
array of combustors defines a second volume different from the
first volume.
2. The rotating detonation combustion system of claim 1, wherein
the first array of combustors and the second array of combustors
are each in alternating adjacent arrangement along the
circumferential direction.
3. The rotating detonation combustion system of claim 2, wherein
the plurality of combustors further defines a third array or more
of combustors in alternating adjacent arrangement along the
circumferential direction with the first array of combustors and
the second array of combustors.
4. The rotating detonation combustion system of claim 1, wherein
the first array of combustors and the second array of combustors
are each in adjacent arrangement along the radial direction, and
the first array of combustors are in adjacent arrangement along the
circumferential direction at a first radius from the longitudinal
centerline, and wherein the second array of combustors are in
adjacent arrangement along the circumferential direction at a
second radius from the longitudinal centerline.
5. The rotating detonation combustion system of claim 4, wherein
the plurality of combustors further defines a third plurality or
more of combustors in adjacent arrangement along the radial
direction with the first array of combustors and the second array
of combustors, and wherein each of the third or more pluralities of
combustors are defined in circumferential arrangement at a third
radius or more from the longitudinal centerline.
6. The rotating detonation combustion system of claim 1, wherein
the first array of combustors defines the first volume configured
to produce a detonation cell width specific to a lowest steady
state operating condition.
7. The rotating detonation combustion system of claim 1, wherein
the second array of combustors defines the second volume configured
to produce a detonation cell width specific to a highest steady
state operating condition.
8. The rotating detonation combustion system of claim 3, wherein
the third array of combustors defines a third volume of the
combustion chamber configured to produce a detonation cell width
specific to an intermediate steady state or transient operating
condition.
9. The rotating detonation combustion system of claim 1, further
comprising a rotating detonation combustor inlet configured to
direct a flow of an oxidizer to one or more of the plurality of
arrays of combustors.
10. The rotating detonation combustion system of claim 9, wherein
the rotating detonation combustor inlet comprises one or more
articulating inlet walls configured to direct the flow of the
oxidizer to an array of combustors.
11. The rotating detonation combustion system of claim 1, further
comprising a rotating detonation combustor outlet configured to
direct a flow of combustion gases from the array of combustors to
an exhaust section.
12. The rotating detonation combustion system of claim 11, wherein
the rotating detonation combustor outlet comprises one or more
articulating outlet walls configured to direct the flow of
combustion gases from a single array of combustors to the exhaust
section.
13. The rotating detonation combustion system of claim 1, wherein
each combustor comprises: a centerbody defining an inner wall of
each combustion chamber; and a fuel injection port defining a fuel
outlet located between the nozzle inlet and the nozzle outlet for
providing fuel to the flow of oxidizer received through the nozzle
inlet.
14. The rotating detonation combustion system of claim 1, wherein
the rotating detonation combustion system is configured to provide
a fuel sequentially to the plurality of arrays of combustors.
15. A propulsion system defining a radial direction, a longitudinal
direction, and a circumferential direction, wherein a longitudinal
centerline extends along the longitudinal direction, and wherein
the propulsion system defines an upstream end and a downstream end,
the propulsion system comprising: an inlet section at the upstream
end into which an oxidizer flows; a rotating detonation combustion
(RDC) system downstream of the inlet section comprising a plurality
of combustors in adjacent arrangement along the circumferential
direction, wherein each combustor defines a combustor centerline
extended through each combustor, and wherein each combustor
comprises an outer wall defining a combustion chamber and a
combustion inlet, wherein each combustion chamber is defined by an
annular gap and a combustion chamber length together defining a
volume of each combustion chamber, and further wherein each
combustor defines a plurality of nozzle assemblies each disposed at
the combustion inlet in adjacent arrangement around each combustor
centerline, and wherein each nozzle assembly defines a nozzle wall
extended along a lengthwise direction, a nozzle inlet, a nozzle
outlet, and a throat therebetween, wherein each nozzle assembly
defines a converging-diverging nozzle, and wherein a first array of
combustors defines a first volume and a second array of combustors
defines a second volume different from the first volume; and an
exhaust section downstream of the RDC system.
16. The propulsion system of claim 15, wherein the RDC system
further defines a third array or more of combustors, and wherein
the third array of combustors each define a third volume of the
combustion chamber of each combustor, and wherein the third array
is configured to produce a detonation cell width specific to an
intermediate steady state or transient operating condition.
17. The rotating detonation combustion system of claim 15, wherein
the first array of combustors defines the first volume configured
to produce a detonation cell width specific to a lowest steady
state operating condition.
18. The rotating detonation combustion system of claim 15, wherein
the second array of combustors defines the second volume configured
to produce a detonation cell width specific to a highest steady
state operating condition.
19. The rotating detonation combustion system of claim 15, wherein
the RDC system further comprises a rotating detonation combustor
inlet configured to direct a flow of an oxidizer from the inlet
section to one or more of the plurality of arrays of
combustors.
20. The rotating detonation combustion system of claim 15, wherein
the RDC system further comprises a rotating detonation combustor
outlet configured to direct a flow of combustion gases from the
array of combustors to the exhaust section.
Description
FIELD
[0001] The present subject matter relates generally to a system of
continuous detonation in a propulsion system.
BACKGROUND
[0002] Many propulsion systems, such as gas turbine engines, are
based on the Brayton Cycle, where air is compressed adiabatically,
heat is added at constant pressure, the resulting hot gas is
expanded in a turbine, and heat is rejected at constant pressure.
The energy above that required to drive the compression system is
then available for propulsion or other work. Such propulsion
systems generally rely upon deflagrative combustion to burn a
fuel/air mixture and produce combustion gas products which travel
at relatively slow rates and constant pressure within a combustion
chamber. While engines based on the Brayton Cycle have reached a
high level of thermodynamic efficiency by steady improvements in
component efficiencies and increases in pressure ratio and peak
temperature, further improvements are welcomed nonetheless.
[0003] Accordingly, improvements in engine efficiency have been
sought by modifying the engine architecture such that the
combustion occurs as a detonation in either a continuous or pulsed
mode. The pulsed mode design involves one or more detonation tubes,
whereas the continuous mode is based on a geometry, typically an
annulus, within which single or multiple detonation waves spin. For
both types of modes, high energy ignition detonates a fuel/air
mixture that transitions into a detonation wave (i.e., a fast
moving shock wave closely coupled to the reaction zone). The
detonation wave travels in a Mach number range greater than the
speed of sound (e.g., Mach 4 to 8) with respect to the speed of
sound of the reactants. The products of combustion follow the
detonation wave at the speed of sound relative to the detonation
wave and at significantly elevated pressure. Such combustion
products may then exit through a nozzle to produce thrust or rotate
a turbine.
[0004] However, propulsion systems, and rotating detonation
combustion systems specifically, are generally designed or
optimized to a specific operating condition or design point (e.g.,
an aero design point) at which the system is most efficient or
operable. Outside or beyond such design points, a rotating
detonation combustion system may be unacceptably inefficient or
inoperable, such as the cell size for a fixed stoichiometry
changing by approximately a factor of 20 across a range of
pressures and temperatures (e.g., from a lowest operating condition
to a highest operating condition), thereby limiting applications of
rotating detonation combustion systems, or offsetting efficiencies
of rotating detonation combustion systems at certain design points
by excessive inefficiencies off design point.
[0005] Therefore, there is a need for a propulsion system and
rotating detonation combustion system that provides efficiency and
operability across a plurality of operating conditions.
BRIEF DESCRIPTION
[0006] Aspects and advantages of the invention will be set forth in
part in the following description, or may be obvious from the
description, or may be learned through practice of the
invention.
[0007] The present disclosure is directed to a rotating detonation
combustion system for a propulsion system. The rotating detonation
combustion system defines a radial direction, a circumferential
direction, and a longitudinal centerline in common with the
propulsion system extended along a longitudinal direction. The
rotating detonation combustion system includes a plurality of
combustors in adjacent arrangement along the circumferential
direction. Each combustor defines a combustor centerline extended
through each combustor, and each combustor comprises an outer wall
defining a combustion chamber and a combustion inlet. Each
combustion chamber is defined by an annular gap and a combustion
chamber length together defining a volume of each combustion
chamber. Each combustor defines a plurality of nozzle assemblies
each disposed at the combustion inlet in adjacent arrangement
around each combustor centerline. Each nozzle assembly defines a
nozzle wall extended along a lengthwise direction, a nozzle inlet,
a nozzle outlet, and a throat therebetween, and each nozzle
assembly defines a converging-diverging nozzle. A first array of
combustors defines a first volume and a second array of combustors
defines a second volume different from the first volume.
[0008] In various embodiments of the rotating detonation combustion
system, the first array of combustors and the second array of
combustors are each in alternating adjacent arrangement along the
circumferential direction. In one embodiment, the plurality of
combustors further defines a third array or more of combustors in
alternating adjacent arrangement along the circumferential
direction with the first array of combustors and the second array
of combustors.
[0009] In still various embodiments, the first array of combustors
and the second array of combustors are each in adjacent arrangement
along the radial direction, and the first array of combustors are
in adjacent arrangement along the circumferential direction at a
first radius from the longitudinal centerline, and wherein the
second array of combustors are in adjacent arrangement along the
circumferential direction at a second radius from the longitudinal
centerline. In one embodiment, the plurality of combustors further
defines a third plurality or more of combustors in adjacent
arrangement along the radial direction with the first array of
combustors and the second array of combustors, in which each of the
third or more pluralities of combustors are defined in
circumferential arrangement at a third radius or more from the
longitudinal centerline.
[0010] In one embodiment, the first array of combustors defines the
first volume configured to produce a detonation cell width specific
to a lowest steady state operating condition.
[0011] In another embodiment, the second array of combustors
defines the second volume configured to produce a detonation cell
width specific to a highest steady state operating condition.
[0012] In still another embodiment, the third array of combustors
defines a third volume of the combustion chamber configured to
produce a detonation cell width specific to an intermediate steady
state or transient operating condition.
[0013] In various embodiments, the rotating detonation combustion
system further includes a rotating detonation combustor inlet
configured to direct a flow of an oxidizer to one or more of the
plurality of arrays of combustors. In one embodiment, the rotating
detonation combustor inlet comprises one or more articulating inlet
walls configured to direct the flow of the oxidizer to an array of
combustors.
[0014] In still various embodiments, the rotating detonation
combustion system further includes a rotating detonation combustor
outlet configured to direct a flow of combustion gases from the
array of combustors to an exhaust section. In one embodiment, the
rotating detonation combustor outlet comprises one or more
articulating outlet walls configured to direct the flow of
combustion gases from a single array of combustors to the exhaust
section.
[0015] In another embodiment, each combustor includes a centerbody
defining an inner wall of each combustion chamber, and a fuel
injection port defining a fuel outlet located between the nozzle
inlet and the nozzle outlet for providing fuel to the flow of
oxidizer received through the nozzle inlet.
[0016] In yet another embodiment, the rotating detonation
combustion system is configured to provide a fuel sequentially to
the plurality of arrays of combustors.
[0017] The present disclosure is further directed to a propulsion
system defining a radial direction, a longitudinal direction, and a
circumferential direction, wherein a longitudinal centerline
extends along the longitudinal direction, and an upstream end and a
downstream end. The propulsion system includes an inlet section at
the upstream end into which an oxidizer flows. The propulsion
system further includes the rotating detonation combustion system
downstream of the inlet section, and an exhaust section downstream
of the rotating detonation combustion system.
[0018] In one embodiment of the propulsion system, the rotating
detonation combustion system further defines a third array or more
of combustors, in which the third array of combustors each define a
third volume of the combustion chamber of each combustor, and the
third array is configured to produce a detonation cell width
specific to an intermediate steady state or transient operating
condition.
[0019] In another embodiment of the propulsion system, the first
array of combustors of the rotating detonation combustion system
defines the first volume configured to produce a detonation cell
width specific to a lowest steady state operating condition. In yet
another embodiment, the second array of combustors defines the
second volume configured to produce a detonation cell width
specific to a highest steady state operating condition.
[0020] In still other embodiments of the propulsion system, the
rotating detonation combustion system further includes a rotating
detonation combustor inlet configured to direct a flow of an
oxidizer from the inlet section to one or more of the plurality of
arrays of combustors.
[0021] In still yet another embodiment of the propulsion system,
the rotating detonation combustion system further includes a
rotating detonation combustor outlet configured to direct a flow of
combustion gases from the array of combustors to the exhaust
section.
[0022] These and other features, aspects and advantages of the
present invention will become better understood with reference to
the following description and appended claims. The accompanying
drawings, which are incorporated in and constitute a part of this
specification, illustrate embodiments of the invention and,
together with the description, serve to explain the principles of
the invention.
BRIEF DESCRIPTION OF THE DRAWINGS
[0023] A full and enabling disclosure of the present invention,
including the best mode thereof, directed to one of ordinary skill
in the art, is set forth in the specification, which makes
reference to the appended figures, in which:
[0024] FIG. 1 is a schematic view of a propulsion system in
accordance with an exemplary embodiment of the present
disclosure;
[0025] FIG. 2 is a cross-sectional circumferential view of a
rotating detonation combustion system in accordance with an
exemplary embodiment of the present disclosure;
[0026] FIG. 3 is a cross-sectional circumferential view of a
rotating detonation combustion system in accordance with another
exemplary embodiment of the present disclosure;
[0027] FIG. 4 is an axial view of the exemplary rotating detonation
combustion system of FIGS. 1-3;
[0028] FIG. 5 is a cross-sectional circumferential view of a
rotating detonation combustion system in accordance with yet
another exemplary embodiment of the present disclosure;
[0029] FIG. 6 is an axial view of the exemplary rotating detonation
combustion system of FIG. 5;
[0030] FIG. 7 is a close-up, axial cross-sectional view of a nozzle
of the exemplary rotating detonation combustion system of FIGS. 1-3
in accordance with an exemplary embodiment of the present
disclosure;
[0031] FIG. 8 is a perspective view of a single combustion chamber
of the exemplary rotating detonation combustion system of FIGS.
1-3;
[0032] FIG. 9 is a cross-sectional circumferential view of an
exemplary embodiment of a single combustor of the rotating
detonation combustion system of FIGS. 1-6;
[0033] FIG. 10 is a cross-sectional circumferential view of another
exemplary embodiment of a single combustor of the rotating
detonation combustion system of FIGS. 1-6;
[0034] FIG. 11 is an axial cross-sectional view of an exemplary
embodiment of the rotating detonation combustion system shown in
FIGS. 1-8;
[0035] FIG. 12 is an axial cross-sectional view of another
exemplary embodiment of the rotating detonation combustion system
shown in FIGS. 1-8;
[0036] FIG. 13 is an axial cross-sectional view of yet another
exemplary embodiment of the rotating detonation combustion system
shown in FIGS. 1-8; and
[0037] FIG. 14 is an axial cross-sectional view of still another
exemplary embodiment of the rotating detonation combustion system
shown in FIGS. 1-8;
DETAILED DESCRIPTION
[0038] Reference will now be made in detail to present embodiments
of the invention, one or more examples of which are illustrated in
the accompanying drawings. The detailed description uses numerical
and letter designations to refer to features in the drawings. Like
or similar designations in the drawings and description have been
used to refer to like or similar parts of the invention.
[0039] As used herein, the terms "first", "second", and "third" may
be used interchangeably to distinguish one component from another
and are not intended to signify location or importance of the
individual components.
[0040] The terms "forward" and "aft" refer to relative positions
within a propulsion system or vehicle, and refer to the normal
operational attitude of the propulsion system or vehicle. For
example, with regard to a propulsion system, forward refers to a
position closer to a propulsion system inlet and aft refers to a
position closer to a propulsion system nozzle or exhaust.
[0041] The terms "upstream" and "downstream" refer to the relative
direction with respect to fluid flow in a fluid pathway. For
example, "upstream" refers to the direction from which the fluid
flows, and "downstream" refers to the direction to which the fluid
flows.
[0042] The singular forms "a", "an", and "the" include plural
references unless the context clearly dictates otherwise.
[0043] Approximating language, as used herein throughout the
specification and claims, is applied to modify any quantitative
representation that could permissibly vary without resulting in a
change in the basic function to which it is related. Accordingly, a
value modified by a term or terms, such as "about",
"approximately", and "substantially", are not to be limited to the
precise value specified. In at least some instances, the
approximating language may correspond to the precision of an
instrument for measuring the value, or the precision of the methods
or machines for constructing or manufacturing the components and/or
systems. For example, the approximating language may refer to being
within a 10 percent margin.
[0044] Here and throughout the specification and claims, range
limitations are combined and interchanged, such ranges are
identified and include all the sub-ranges contained therein unless
context or language indicates otherwise. For example, all ranges
disclosed herein are inclusive of the endpoints, and the endpoints
are independently combinable with each other.
[0045] Embodiments of a propulsion system and rotating detonation
combustion (RDC) system that provide efficiency and operability
across a plurality of operating conditions are generally provided.
The embodiments generally provided herein enable a desired or
optimum detonation cell width and/or cell quantity in one or more
combustors of the rotating detonation combustion system across a
plurality of operating conditions defining a plurality of pressures
and temperatures at the RDC system. The embodiments generally
provided herein include a plurality of combustors configured to
provide a desired or optimum detonation cell width and/or quantity
for a plurality of operating conditions of the RDC system and the
propulsion system. As such, the RDC system and propulsion system
may generally provide improved specific fuel consumption and fuel
burn while also improving combustion stability, emissions,
lean-blowout mitigation, and propulsion system operability.
[0046] Referring now to the figures, FIG. 1 depicts a propulsion
system including a rotating detonation combustion system 100 (an
"RDC system") in accordance with an exemplary embodiment of the
present disclosure. For the embodiment of FIG. 1, the propulsion
system 102 generally includes an inlet section 104 and an exhaust
section 106, with the RDC system 100 located downstream of the
inlet section 104 and upstream of the exhaust section 106. The
propulsion system 102 defines a longitudinal direction L and a
radial direction R. A longitudinal centerline 116 extends through
the propulsion system 102 along the longitudinal direction L and is
provided for illustrative purposes. In various embodiments, the
propulsion system 102 defines a gas turbine engine, a ramjet, or
other propulsion system including a fuel-oxidizer burner producing
combustion products that provide propulsive thrust or mechanical
energy output. In an embodiment of the propulsion system 102
defining a gas turbine engine, the inlet section 104 includes a
compressor section defining one or more compressors generating a
flow of oxidizer to the RDC system 100. The inlet section 104 may
generally guide a flow of the oxidizer 131 to the RDC system 100.
The inlet section 104 may further compress the oxidizer before it
enters the RDC system 100. The inlet section 104 defining a
compressor section may include one or more alternating stages of
rotating compressor airfoils. In other embodiments, the inlet
section 104 may generally define a decreasing cross sectional area
from an upstream end to a downstream end proximate to the RDC
system 100.
[0047] As will be discussed in further detail below, at least a
portion of the flow of oxidizer is mixed with a fuel 163 (shown in
FIG. 5) to generate combustion products 138. The combustion
products 138 flow downstream to the exhaust section 106. In various
embodiments, the exhaust section 106 may generally define an
increasing cross sectional area from an upstream end proximate to
the RDC system 100 to a downstream end of the propulsion system
102. Expansion of the combustion products 138 generally provides
thrust that propels the apparatus to which the propulsion system
102 is attached, or provides mechanical energy to one or more
turbines further coupled to a fan section, a generator, or both.
Thus, the exhaust section 106 may further define a turbine section
of a gas turbine engine including one or more alternating rows or
stages of rotating turbine airfoils. The combustion products 138
may flow from the exhaust section 106 through, e.g., an exhaust
nozzle 135 to generate thrust for the propulsion system 102.
[0048] As will be appreciated, in various embodiments of the
propulsion system 102 defining a gas turbine engine, rotation of
the turbine(s) within the exhaust section 106 generated by the
combustion products is transferred through one or more shafts or
spools 110 to drive the compressor(s) within the inlet section 104.
In various embodiments, the inlet section 104 may further define a
fan section, such as for a turbofan engine configuration, such as
to propel air across a bypass flowpath outside of the RDC system
100 and exhaust section 106. The combustion products may then flow
from the exhaust section 106 through, e.g., an exhaust nozzle 135
to generate thrust for the propulsion system 102.
[0049] It will be appreciated that the propulsion system 102
depicted schematically in FIG. 1 is provided by way of example
only. In certain exemplary embodiments, the propulsion system 102
may include any suitable number of compressors within the inlet
section 104, any suitable number of turbines within the exhaust
section 106, and further may include any number of shafts or spools
110 appropriate for mechanically linking the compressor(s),
turbine(s), and/or fans. Similarly, in other exemplary embodiments,
the propulsion system 102 may include any suitable fan section,
with a fan thereof being driven by the exhaust section 106 in any
suitable manner. For example, in certain embodiments, the fan may
be directly linked to a turbine within the exhaust section 106, or
alternatively, may be driven by a turbine within the exhaust
section 106 across a reduction gearbox. Additionally, the fan may
be a variable pitch fan, a fixed pitch fan, a ducted fan (i.e., the
propulsion system 102 may include an outer nacelle surrounding the
fan section), an un-ducted fan, or may have any other suitable
configuration.
[0050] Moreover, it should also be appreciated that the RDC system
100 may further be incorporated into any other suitable
aeronautical propulsion system, such as a turboshaft engine, a
turboprop engine, a turbojet engine, a ramjet engine, a scramjet
engine, etc. Further, in certain embodiments, the RDC system 100
may be incorporated into a non-aeronautical propulsion system, such
as a land-based or marine-based power generation system. Further
still, in certain embodiments, the RDC system 100 may be
incorporated into any other suitable propulsion system, such as a
rocket or missile engine. With one or more of the latter
embodiments, the propulsion system may not include a compressor in
the inlet section 104 or a turbine in the exhaust section 106.
[0051] Referring now to FIGS. 2-3, cross sectional views providing
a circumferential view of exemplary embodiments of the RDC system
100 are generally provided. The RDC system 100 includes a plurality
of combustors 112 disposed in adjacent arrangement along a
circumferential direction C around the longitudinal centerline 116
of the propulsion system 102. Each combustor 112 defines a
combustor centerline 115 in which each combustor centerline 115
disposes each combustor 112 in circumferential arrangement around
the longitudinal centerline 116 of the propulsion system 102. Each
combustor 112 includes an outer wall 118 defining a generally
cylindrical walled enclosure. As a non-limiting example, the outer
wall 118 defines each combustor 112 as a can combustor.
[0052] In one embodiment of the RDC system 100, such as generally
provided in FIG. 2, the plurality of combustors 112 are disposed in
adjacent circumferential arrangement generally at an approximately
common radius from the longitudinal centerline 116. For example,
each combustor centerline 115 of the plurality of combustors 112 is
arranged at an approximately equal radius from the longitudinal
centerline 116. The various embodiments of the plurality of
combustors 112, as further described below as different arrays
(e.g., first array 166, second array 168, third array 170, Nth
array, etc.) are each alternating and disposed in adjacent
circumferential arrangement from a first radius from the
longitudinal centerline 116.
[0053] In another embodiment of the RDC system 100, such as
generally provided in FIG. 3, the plurality of combustors 112 are
each disposed in adjacent circumferential arrangement and
furthermore in radial arrangement. For example, a first array 166
is disposed at a first radius from the longitudinal centerline 116;
a second array 168 is disposed at a second radius from the
longitudinal centerline 116 different from the first radius; and a
third array 170 is disposed at a third radius from the longitudinal
centerline 116 different from the first and second radii. In other
embodiments, an Nth array may be disposed at an Nth radius from the
longitudinal centerline 116 different from the other radii (e.g.,
first radius, second radius . . . Nth radius, etc.).
[0054] Referring to FIG. 4, within each walled enclosure defined by
the outer wall 118 is defined a combustion chamber 122, a
combustion chamber inlet 124 proximate to a nozzle assembly 128,
and a combustion chamber outlet 126. The nozzle assembly 128
provides a flow mixture of oxidizer and fuel to the combustion
chamber 122, wherein such mixture is combusted/detonated to
generate the combustion products therein, and more specifically a
detonation wave 130 as will be explained in greater detail below.
The combustion products exit through the combustion chamber outlet
126. Each combustion chamber 122 defines a width or annular gap 121
and a combustion chamber length 123. The annular gap 121 and the
combustion chamber length 123 together define a volume of each
combustion chamber 122. The annular gap 121 and the combustion
chamber length 123 are each variables at least partly determinate
of a desired or optimal operating condition of the RDC system 100
and propulsion system 102. The annular gap 121 is defined as
generally encompassing an area at which a mixture of fuel and
oxidizer 195 (such as the fuel/oxidizer mixture 132 shown in FIG.
8) is present within the combustion chamber 122. In various
embodiments, such as generally provided in FIGS. 2, 3, and 9, the
annular gap 121 extends generally from the outer wall 118 to a
generally cylindrical inner wall 120. In other embodiments, such as
generally provided in FIGS. 4 and 10, the annular gap 121 is
defined from the outer wall 118 to an inner diameter at which fuel
is present in the combustion chamber 122, such as schematically
shown by lines 107.
[0055] Referring now to FIGS. 2-4, the RDC system 100 defines a
plurality of combustors 112 defining a plurality of volumes of
combustion chambers 122. The RDC system 100 may generally define a
first volume 201 corresponding to a first array 166 of combustors
112 and a second volume 202 corresponding to a second array 168 of
combustors 112. In various embodiments, the RDC system 100 may
further include a third volume 203 corresponding to a third array
170 of combustors 112, or a fourth, fifth, Nth, etc. volume
corresponding to a fourth, first, Nth, etc. array of combustors
112. Each volume 201, 202, 203 is defined generally within the
outer wall 118. For example, the outer wall 118 defines a diameter
within which the combustion chamber 122 is defined, such as
generally provided in FIG. 10. In various embodiments, such as
generally provided in FIGS. 2, 3, and 9, the volume 201, 202, 203
of the combustion chamber 122 may further be defined between the
outer wall 118 and the inner wall 120.
[0056] Each array defines each volume with an annular gap 121, a
combustion chamber length 123, or both, different from each other
array. Each volume (e.g., first volume 201, second volume 202,
third volume 203, etc.) configures the annular gap 121, the
combustion chamber length 123, or both to produce a desired or
optimal quantity of detonation cells within each combustion chamber
122 based on the annular gap 121, the combustion chamber length
123, or both of each array 166, 168, 170 of combustors 112 and
further based on a plurality of design points or operating
conditions specific to each array 166, 168, 170.
[0057] For example, in one embodiment, the first array 166 of
combustors 112 defines the first volume 201 (i.e., the annular gap
121, the combustion chamber length 123, or both) based on a desired
quantity of detonation cells or cell size corresponding to pressure
and temperatures conditions for a lowest steady state operating
condition of the propulsion system 102 (e.g., lowest pressure,
lowest temperature, or both at the RDC system 100 above initial
light-off or startup). In an embodiment in which the propulsion
system 102 defines a propulsive gas turbine engine, the lowest
steady state operating condition may define a ground idle
condition. In various embodiments, the first array 166 of
combustors 112 may define the first volume 201 as a maximum volume
relative to the plurality of combustors 112 defining a plurality of
volumes.
[0058] As another example, in another embodiment, the third array
170 of combustors 112 defines the third volume 203 (i.e., the
annular gap 121, the combustion chamber length 123, or both) based
on a desired quantity of detonation cells or cell size
corresponding to pressure and temperatures conditions for a highest
steady state operating condition of the propulsion system 102
(e.g., highest pressure, highest temperature, or both at the RDC
system 100). In an embodiment in which the propulsion system 102
defines a propulsive gas turbine engine, the highest steady state
operating condition may define a maximum takeoff condition. In
various embodiments, the third array 170 of combustors 112 may
define the third volume 203 as a minimum volume relative to the
plurality of combustors 112 defining a plurality of volumes.
[0059] As yet another example, in another embodiment, the second
array 168 of combustors 112 defines the second volume 202 (i.e.,
the annular gap 121, the combustion chamber length 123, or both)
based on a desired quantity of detonation cells or cell size
corresponding to pressure and temperatures conditions for an
intermediate steady state or transient operating condition of the
propulsion system 102. In an embodiment in which the propulsion
system 102 defines a propulsive gas turbine engine, the
intermediate steady state or transient operating condition may
define one or more conditions greater than ground idle and less
than a maximum takeoff condition, such as, but not limited to,
cruise, climb, flight idle, or approach. In various embodiments,
the second array 170 of combustors 112 may define the second volume
202 greater than the first volume 201 and less than the third
volume 203. In still another example, the second array 168 may be
configured to an intermediate steady state operating condition
corresponding to a cruise condition for an aero gas turbine engine.
The intermediate steady state operating condition, such as a cruise
condition, may be based further on one or more of a duration of
time expected at the intermediate steady state operating condition,
altitude, flow rate of oxidizer into the RDC system 100, pressure
and/or temperature of the oxidizer and/or fuel, or combinations
thereof
[0060] Referring now to FIG. 5, a cross sectional view providing a
circumferential view of another exemplary embodiment of the RDC
system 100 is generally provided. The embodiment provided in FIG. 5
may be configured substantially similarly as shown and described in
regard to FIGS. 1-4. However, in FIG. 5, each combustor 112 is
disposed in adjacent radial arrangement from the longitudinal
centerline 116. Each combustor 112 defines an array (e.g., array
166, 168, 170) disposed in generally concentric arrangement around
the longitudinal centerline 116 (e.g., defining concentric annular
combustor rings). As discussed in regard to FIGS. 2-4, each array
defines a volume configured to a corresponding operating condition
of the propulsion system 102.
[0061] Referring now to FIG. 6, a schematic cross sectional view is
generally provided of the exemplary embodiment of the RDC system
100 shown in FIG. 5. Referring to FIGS. 5-6, each combustor 112 is
defined by at least two of an outer wall 118, an intermediate wall
113, and an inner wall 120 extended along the longitudinal
direction L from the nozzle assembly 128 and combustion inlet 124
to the combustion outlet 126. Each pair of walls 118, 113, 120
defines a combustion chamber 122 therebetween. For example, the
outer wall 118 and one of the intermediate walls 113 defines the
first array 166 defining the first volume 201 of the combustion
chamber 122. A pair of intermediate walls 113 defines the second
array 168 defining the second volume 202 of the combustion chamber
122. The inner wall 120 and one of the intermediate walls 113
defines the third array 170 defining the third volume 203. It
should be appreciated that a plurality of pairs of intermediate
walls 113 may further define additional arrays each defining
additional volumes (e.g. fourth array, fifth array . . . Nth
array). It should further be appreciated that the annular gap 121
is defined specific to each pair of the outer wall 118, the
intermediate walls 113, and the inner wall 120 each defining a
volume 201, 202, 203 of the combustion chamber 122.
[0062] Referring now to FIG. 7, a schematic cross sectional view is
provided of a portion of an exemplary nozzle assembly 128 of the
RDC system 100 as may be incorporated into the exemplary
embodiments of FIGS. 2-4. As shown, each nozzle 140 of the RDC
system 100 generally defines a nozzle centerline 117 and a radial
direction RR extended therefrom relative to the nozzle centerline
117. The nozzle 140 extends along the lengthwise direction 142
between a nozzle inlet 144 and a nozzle outlet 146, and further
defines a nozzle flowpath 148 extending from the nozzle inlet 144
to the nozzle outlet 146.
[0063] In various embodiments, the nozzle 140 includes a nozzle
wall 150 defining the nozzle flowpath 148, such as shown in FIG. 4.
In one embodiment, the nozzle wall 150 extends circumferentially
around each nozzle centerline 117. As discussed further below in
regard to FIGS. 9-10, in embodiments such as generally provided in
FIGS. 2-3, each nozzle 140 and its corresponding nozzle centerline
117 is disposed in adjacent circumferential arrangement around the
combustor centerline 115 of each combustor 112. As such, the nozzle
wall 150 may define an annular wall around each nozzle centerline
117. However, in other embodiments, such as generally provided in
FIGS. 5-6, each nozzle 140 is defined around the longitudinal
centerline 116 of the propulsion system 102. As such, the nozzle
wall 150 may define an annular wall around the longitudinal
centerline 116.
[0064] In still various embodiments, the nozzle wall 150 is a
continuous nozzle wall extending from the nozzle inlet 144 to the
nozzle outlet 146. However, in other embodiments, the nozzle wall
150 may have any other suitable configuration. In various
embodiments, the nozzle 140 defines a converging-diverging nozzle,
in which the nozzle wall 150 decreases the nozzle flowpath area
from approximately the nozzle inlet 144 to approximately a throat
152 between the nozzle inlet 144 and nozzle outlet 146, and in
which the nozzle wall 150 increases the nozzle flowpath area from
approximately the throat 152 to approximately the nozzle outlet
146.
[0065] Referring still to FIG. 7, the nozzle inlet 144 is
configured to receive a flow of oxidizer during operation of the
RDC system 100 and provide such flow oxidizer through/along the
nozzle flowpath 148. The flow of oxidizer may be a flow of air,
oxygen, etc. More specifically, when the nozzle 140 of the nozzle
assembly 128 is incorporated into the RDC system 100 of the
propulsion system 102 of FIG. 1, the flow of oxidizer will be a
flow of compressed air from the inlet section 104.
[0066] The nozzle 140, or more specifically the nozzle wall 150,
further defines the throat 152 between the nozzle inlet 144 and the
nozzle outlet 146, i.e., downstream of the nozzle inlet 144 and
upstream of the nozzle outlet 146. As used herein, the term
"throat", with respect to the nozzle 140, refers to the point
within the nozzle flowpath 148 having the smallest cross-sectional
area. Additionally, as used herein, the term "cross-sectional
area", such as a cross-sectional area of the throat 152, refers to
an area within the nozzle flowpath 148 at a cross-section measured
along the radial direction RR at the respective location along the
nozzle flowpath 148.
[0067] In various embodiments, the nozzle 140 may be referred to as
a converging-diverging nozzle. Further, for the embodiment
depicted, the throat 152 is positioned closer to the nozzle inlet
144 than the nozzle outlet 146 along the lengthwise direction 142
of the nozzle 140. More specifically, as is depicted, the nozzle
140 defines a length 160 along the lengthwise direction 142. The
throat 152 for the exemplary nozzle 140 depicted is positioned in a
forward, or upstream, portion of the length 160 of the nozzle 140.
More specifically, still, for the embodiment depicted the throat
152 of the exemplary nozzle 140 depicted is positioned
approximately between the forward ten percent and fifty percent of
the length 160 of the nozzle 140 along the lengthwise direction
142, such as approximately between the forward twenty percent and
forty percent of the length 160 of the nozzle 140 along the
lengthwise direction 142.
[0068] A nozzle 140 having such a configuration may provide for a
substantially subsonic flow through the nozzle flowpath 148. For
example, the flow from the nozzle inlet 144 to the throat 152
(i.e., a converging section 159 of the nozzle 140) may define an
airflow speed below Mach 1. The flow through the throat 152 may
define an airflow speed less than Mach 1, but approaching Mach 1,
such as within about ten percent of Mach 1, such as within about
five percent of Mach 1.
[0069] Additionally, the flow from the throat 152 to the nozzle
outlet 146 (i.e., a diverging section 161 of the nozzle 140) may
again define an airflow speed below Mach 1 and less than the
airflow speed through the throat 152. In other embodiments, the
airflow speed may be Mach 1 downstream of the throat 152. For
example, a small region downstream of the throat 152 may define an
airflow speed at or above Mach 1 before defining a weak normal
shock to less than Mach 1.
[0070] Referring still to FIG. 7, the RDC system 100 further
includes a fuel injection port 162. The fuel injection port 162
defines a fuel outlet 164 in fluid communication with the nozzle
flowpath 148 and located between the nozzle inlet 144 and the
nozzle outlet 146 for providing fuel to the flow of oxidizer
received through the nozzle inlet 144. More specifically, in
various embodiments, the fuel outlet 164 of the fuel injection port
162 is positioned within a buffer distance from the throat 152 of
the nozzle 140 along the lengthwise direction 142 of the nozzle 140
(with the buffer distance being a distance equal to ten percent of
the length 160 of the nozzle 140 along the lengthwise direction
142). More particularly, for the embodiment depicted, the fuel
outlet 164 of the fuel injection port 162 is positioned at the
throat 152 of the nozzle 140, or downstream of the throat 152 of
the nozzle 140 along the lengthwise direction 142 of the nozzle
140. More specifically still, for the embodiment depicted, the fuel
outlet 164 of the fuel injection port 162 is positioned at the
throat 152 of the nozzle 140. It will be appreciated, that as used
herein, the term "at the throat of the nozzle" refers to including
at least a portion of the component or feature positioned at a
location within the nozzle flowpath 148 defining the smallest
cross-sectional area (i.e., defining the throat 152). Notably, for
the embodiment of FIG. 5, the throat 152 of the exemplary nozzle
140 depicted is not a single point along the lengthwise direction
142, and instead extends for a distance along the lengthwise
direction 142. For the purposes of measuring locations of features
or parts relative to the throat 152, the measurement may be taken
from anywhere within the nozzle flowpath 148 defining the throat
152. Notably, although the fuel injection port 162 is depicted as
including two outlets 164 in radially adjacent arrangement, it
should be understood that a plurality of fuel injection ports 162
may be distributed along the circumferential direction (relative to
the nozzle centerline 117) along the annulus of the nozzle 140.
[0071] The fuel provided through the fuel injection port 162 may be
any suitable fuel, such as a hydrocarbon-based fuel, for mixing
with the flow of oxidizer. More specifically, for the embodiment
depicted the fuel injection port 162 is a liquid fuel injection
port configured to provide a liquid fuel to the nozzle flowpath
148, such as a liquid jet fuel. However, in other exemplary
embodiments, the fuel may be a gaseous fuel, or a mixture of a
liquid fuel and gaseous fuel, or a mixture of a liquid fuel and a
non-fuel gas, or any other suitable fuel, or combinations
thereof
[0072] Accordingly, for the embodiment depicted, positioning the
fuel outlet 164 of the fuel injection port 162 in accordance with
the description above may allow for the liquid fuel provided
through the outlet 164 of the fuel injection port 162 to
substantially completely atomize within the flow of oxidizer
provided through the nozzle inlet 144 of the nozzle 140. Such may
provide for a more complete mixing of the fuel within the flow of
oxidizer, providing for a more complete and stable combustion
within the combustion chamber 122.
[0073] Furthermore, for the embodiment depicted, the fuel injection
port 162 is integrated into the nozzle 140. More specifically, for
the embodiment depicted, the fuel injection port 162 extends
through, and may be at least partially defined by, or positioned
within, an opening extending through the nozzle wall 150 of the
nozzle 140. Additionally, for the embodiment, the fuel injection
port 162 further includes a plurality of fuel injection ports 162,
with each fuel injection port 162 defining an outlet 164. In
various embodiments, the plurality of fuel injection ports 162,
each defining the outlet 164, are arranged along the
circumferential direction around the longitudinal centerline 116.
The plurality of fuel injection ports 162 may be arranged in
symmetric or asymmetric arrangement around the longitudinal
centerline 116.
[0074] Each of the one or more fuel injection ports 162 may be
fluidly connected to a fuel source, such as a fuel tank, through
one or more fuel lines for supplying the fuel to the fuel injection
ports 162 (not shown). Additionally, it should be appreciated, that
in other exemplary embodiments, the fuel injection port 162 may not
be integrated into the nozzle 140. With such an exemplary
embodiment, the RDC system 100 may instead include a fuel injection
port having a separate structure extending, e.g., through the
nozzle inlet 144 and nozzle flowpath 148. Such a fuel injection
port may further define a fuel outlet positioned in the nozzle
flowpath 148 between the nozzle inlet 144 and the nozzle outlet 146
for providing fuel to the flow of oxidizer received through the
nozzle inlet 144.
[0075] A nozzle 140 in accordance with one or more of the exemplary
embodiments described herein may allow for a relatively low
pressure drop from the nozzle inlet 144 to the nozzle outlet 146
and into the combustion chamber 122. For example, in certain
exemplary embodiments, a nozzle 140 in accordance with one or more
of the exemplary embodiments described herein may provide for a
pressure drop of less than about twenty percent. For example, in
certain exemplary embodiments the nozzle 140 may provide for a
pressure drop less than about twenty-five percent, such as between
about one percent and about fifteen percent, such as between about
one percent and about ten percent, such as between about one
percent and about eight percent, such as between about one percent
and about six percent. It should be appreciated, that as used
herein, the term "pressure drop" refers to a pressure difference
between a flow at the nozzle outlet 146 and at the nozzle inlet
144, as a percentage of the pressure of the flow at the nozzle
inlet 144. Notably, including a nozzle 140 having such a relatively
low pressure drop may generally provide for a more efficient RDC
system 100. In addition, inclusion of a nozzle 140 having a
converging-diverging configuration as is depicted and/or described
herein may prevent or greatly reduce a possibility of the high
pressure fluid (e.g., combustion products) within the region 134
behind the detonation wave 130 from flowing in an upstream
direction, i.e., into the incoming fuel/air mixture flow 132 (see
FIG. 6).
[0076] In various embodiments, each nozzle 140 in the plurality of
nozzles 140 may be configured in accordance with one or more of the
embodiments described above with reference to FIG. 7. Further, in
certain embodiments, each nozzle 140 in the plurality of nozzles
140 may be configured in substantially the same manner, or
alternatively, in other embodiments, one or more of the plurality
of nozzles 140 may include a varied geometry. Furthermore, in still
various embodiments, each nozzle 140 in the plurality of nozzles
140 may be configured in accordance with one or more of the
embodiments described above with reference to FIG. 7. Further, in
certain embodiments, each nozzle 140 in the plurality of nozzles
140 may be configured in substantially the same manner, or
alternatively, in other embodiments, one or more of the plurality
of nozzles 140 may include a varied geometry. For example, the
nozzle wall 150 of each nozzle 140 may define varied
converging-diverging geometries, such as varying angles relative to
the longitudinal centerline 116. In still other embodiments, the
fuel injection ports 162 of each nozzle 140 may define various
areas, volumes, flowpaths, or other flow characteristics relative
to each nozzle 140, or relative to various circumferential
locations within each nozzle 140. In yet other embodiments, the
nozzles 140 may be evenly spaced from one another between the outer
wall 118 and the inner wall 120. In other embodiments, the nozzles
140 may be disposed in uneven arrangement such that one nozzle 140
defines a larger or smaller throat 152 than another nozzle 140.
Still further, although each of the plurality of nozzles 140 is
depicted as including a substantially circular nozzle inlet 144
(and a substantially circular nozzle flowpath 148 along the
respective lengthwise directions 142), in other embodiments, one or
more of the plurality of nozzles 140 instead define any other
suitable cross-sectional shape along a respective lengthwise
direction 142, such as an ovular shape, a polygonal shape, etc.
Similarly, although the converging and diverging sections 159, 161
are depicted as conical, in other exemplary embodiments, one or
both of the sections 159, 161 may be defined by curved walls, or
any other suitable shape. Additionally, the throat 152 of the
nozzle 140 may be a single point along the longitudinal direction
L, as opposed to an elongated cylindrical section.
[0077] Referring briefly to FIG. 8, providing a perspective view of
the combustion chamber 122 (without the nozzle assembly 128 shown),
it will be appreciated that the RDC system 100 generates the
detonation wave 130 during operation. The detonation wave 130
travels in the circumferential direction C relative to the
combustor centerline 115 consuming an incoming fuel/oxidizer
mixture 132 and providing a high pressure region 134 within an
expansion region 136 of the combustion. A burned fuel/oxidizer
mixture 138 (i.e., combustion products) exits the combustion
chamber 122 and is exhausted.
[0078] More particularly, it will be appreciated that the RDC
system 100 is of a detonation-type combustor, deriving energy from
the continuous wave 130 of detonation. For a detonation combustor,
such as the RDC system 100 disclosed herein, the combustion of the
fuel/oxidizer mixture 132 is effectively a detonation as compared
to a burning, as is typical in the traditional deflagration-type
combustors. Accordingly, a main difference between deflagration and
detonation is linked to the mechanism of flame propagation. In
deflagration, the flame propagation is a function of the heat
transfer from a reactive zone to the fresh mixture, generally
through conduction. By contrast, with a detonation combustor, the
detonation is a shock induced flame, which results in the coupling
of a reaction zone and a shockwave. The shockwave compresses and
heats the fresh mixture 132, increasing such mixture 132 above a
self-ignition point. On the other side, energy released by the
combustion contributes to the propagation of the detonation
shockwave 130. Further, with continuous detonation, the detonation
wave 130 propagates around the combustion chamber 122 in a
continuous manner, operating at a relatively high frequency.
Additionally, the detonation wave 130 may be such that an average
pressure inside the combustion chamber 122 is higher than an
average pressure within typical combustion systems (i.e.,
deflagration combustion systems).
[0079] Accordingly, the region 134 behind the detonation wave 130
has very high pressures. As will be appreciated from the discussion
below, the nozzle assembly 128 of the RDC system 100 is designed to
prevent the high pressures within the region 134 behind the
detonation wave 130 from flowing in an upstream direction, i.e.,
into the incoming flow of the fuel/oxidizer mixture 132.
[0080] Referring now to FIGS. 9-10, exemplary embodiments of the
combustors 112 are generally provided from a downstream end viewed
toward upstream. Each embodiment includes a plurality of nozzles
140 each disposed around each nozzle centerline 117. Furthermore,
each nozzle centerline 117 is disposed in circumferential
arrangement around each combustor centerline 115. In the embodiment
provided in FIG. 7, the combustor 117 defines a centerbody 119 that
further defines an inner wall 120 to each combustion chamber 122.
In other words, in the embodiment provided in FIG. 9, each
combustion chamber 122 is defined by the outer wall 118 and the
inner wall 120. The annular gap 121 extends between the inner wall
120 and the outer wall 118, together at least partially defining
the volume of the combustion chamber 122. In other words, the
annular gap 121 is defined by the area between the outer wall 118
and the inner wall 120.
[0081] Referring now to FIGS. 11-14, cross sectional views of at
least a portion of several exemplary embodiments the RDC system 100
are generally provided. The embodiments generally provided include
the plurality of combustors 112 arranged in radial arrangement from
the longitudinal centerline 116 of the propulsion system 102 such
as generally provided in FIG. 3. However, it should be appreciated
that, in other embodiments, each of the embodiments generally
provided in FIGS. 9-12 may be arranged in circumferential
arrangement around the longitudinal centerline 116.
[0082] As generally provided in FIGS. 9-12, the RDC system 100 may
further define a rotating detonation combustor inlet 200 ("RDC
inlet 200") upstream of the plurality of nozzles 140 of each
combustor 112 and a rotating detonation combustor outlet 250 ("RDC
outlet 250") downstream of the combustion chamber 122 of each
combustor 112. The RDC inlet 200 is configured to direct the flow
of oxidizer 131 from the inlet section 104 into each of the arrays
of combustors 112 (e.g., the first array 166, the second array 168,
the third array 170 . . . the Nth array). Furthermore, the RDC
outlet 250 is configured to direct the flow of combustion products
138 from the combustion chamber 122 of each combustor 112 to the
exhaust section 106.
[0083] Referring to FIGS. 12-14, the RDC inlet 200 may further
include an articulating inlet wall 210 configured to direct the
flow of oxidizer 131 to one array of combustors 112 and at least
substantially block the flow of oxidizer 131 from entering the
other arrays of combustors 112. The inlet wall 210 may define a
plurality of vanes, walls, doors, or valves. The inlet wall 210 may
hinge about an axis such as to define a flowpath toward one array
of combustors 112. In one embodiment, such as generally provided in
FIGS. 12-14, the inlet wall 210 hinges or rotates about an axis
defined at a forward or upstream of the inlet wall 210. However, in
various embodiments, the inlet wall 210 may rotate about an axis
defined mid-span or at an aft or downstream end of the inlet wall
210.
[0084] Referring still to FIGS. 12-14, the RDC inlet 200 may
further include an articulating outlet wall 220 configured to
direct the flow of combustion gases 138 from one array of
combustors 112 toward the exhaust section 106. Furthermore, the
outlet wall 220 may prevent a backflow to other arrays of
combustors 112. The outlet wall 220 may define a plurality of
vanes, walls, doors, or valves. The outlet wall 220 may hinge about
an axis such as to define a flowpath toward one array of combustors
112. In one embodiment, such as generally provided in FIGS. 10-12,
the outlet wall 220 hinges or rotates about an axis defined at a
forward or upstream of the outlet wall 220. However, in various
embodiments, the outlet wall 220 may rotate about an axis defined
mid-span or at an aft or downstream end of the outlet wall 220. In
various embodiments, the articulating outer wall 220 directs the
flow of combustion gases 138 from a single array of combustors 112
to the exhaust section 106 based at least on an operating condition
of the RDC system 100 and propulsion system 102, such as further
described below.
[0085] The plurality of inlet walls 210 may be coupled to a common
rail or bracket system coupled to an actuator such that the
plurality of inlet walls 210 disposed circumferentially around the
longitudinal centerline 116 may articulate or rotate in unison. The
plurality of outlet walls 220 may be coupled to a common rail or
bracket system coupled to an actuator such that the plurality of
outlet walls 220 disposed circumferentially around the longitudinal
centerline 116 may articulate or rotate in unison. In still another
embodiment, the plurality of inlet walls 210 and outlet walls 220
may together be configured to articulate or rotate in unison based
at least on an operating condition of the propulsion system
102.
[0086] Referring now to FIGS. 1-14, during operation of the
propulsion system 102, a flow of oxidizer 131 enters through an
inlet 108 of the propulsion system 102 into the inlet section 104.
The inlet section 104 may generally compress the flow of oxidizer
131 before the oxidizer enters the RDC system 100. The RDC inlet
210, as generally provided in FIGS. 10-12, guides or directs the
flow of oxidizer to the arrays of combustors 112 and, more
specifically, the nozzle assembly 128 of each combustor 112. A flow
of fuel 163 (shown in FIG. 5) enters the nozzle flowpath 148 of
each nozzle 140 through the fuel outlet 162. The flow of oxidizer
131 and the fuel 163 together mix (shown schematically by the
fuel-oxidizer mixture 132 shown in FIGS. 7-8) and are detonated in
the combustion chamber 122 of each combustor 112. The combustion
products 138 from the detonation flow downstream to the exhaust
section 106 of the propulsion system 102. The combustion products
138 generate a propulsive force through the exhaust section 106
defining a nozzle structure, a turbine structure, or both.
[0087] More specifically, during an operation of the propulsion
system 102, a flow of fuel 163 is delivered sequentially to one or
more arrays of combustors 112 based at least on an operating
condition of the propulsion system 102. For example, during light
off and a lowest steady state operating condition (e.g., ground
idle), a flow of fuel 163 is delivered to the first array 166 of
combustors 112 to be mixed with the flow of oxidizer 131 and
detonated to yield combustion gases 138. The flow of fuel 163 may
be restricted from delivery to the second array 168 and the third
array 170 by way of independently adjustable fuel manifolds, fuel
lines, or valves (not shown).
[0088] During a transition from a lowest steady state operating
condition to an intermediate or highest steady state operating
condition, the flow of fuel 163 may be delivered to one or more of
the first array 166, the second array 168, and the third array 170
of combustors 112 to be mixed with the flow of oxidizer 131 and
detonated to yield combustion gases 138. At a highest steady state
operating condition (e.g., maximum takeoff) or an intermediate
steady state operating condition (e.g., between ground idle and
maximum takeoff, or cruise condition, etc.), the flow of fuel 163
may be delivered exclusively to the third array 170 or the second
array 168, each configured to a respective operating condition.
[0089] Still further, in various embodiments, during an operation
of the propulsion system 102, the inlet walls 210 and the outlet
walls 220 each articulate or rotate based on the operating
condition. For example, as shown in FIG. 13, the inlet walls 210
and the outlet walls 220 may articulate to direct the flow of
oxidizer 131 to the first array 166 of combustors 112 (e.g.,
defining a first volume 201 optimized for a lowest pressure and
temperature at the combustion chamber 122) during a lowest steady
state operating condition of the propulsion system 102. The outlet
walls 220 articulate to substantially block the second array 168
and the third array 170 to prevent combustion gases 138 from the
first array 166 from backflowing into the second array 168 and
third array 170 of combustors 112. The outer walls 220 may further
substantially block a flow of oxidizer 195 through the second array
168 and the third array 170 from flowing downstream to the exhaust
section 106 while enabling a flow of combustion gases 138 from the
first array 166.
[0090] As another example, as shown in FIG. 14, the inlet walls 210
and the outlet walls 220 may articulate to direct the flow of
oxidizer 131 to the third array 170 of combustors 112 (e.g.,
defining a third volume 203 optimized for a highest pressure and
temperature at the combustion chamber 122) during a highest steady
state operating condition of the propulsion system 102. The outlet
walls 220 articulate to substantially block the first array 166 and
the second array 168 to prevent combustion gases 138 from the third
array 170 from backflowing into the first array 166 and second
array 168 of combustors 112. The outer walls 220 may further
substantially block a flow of oxidizer 195 through the first array
166 and the second array 168 from flowing downstream to the exhaust
section 106 while enabling a flow of combustion gases 138 from the
third array 170.
[0091] As yet another example, as shown in FIG. 12, the inlet walls
210 and the outlet walls 220 may articulate to direct the flow of
oxidizer 131 to the second array 168 of combustors 112 (e.g.,
defining a second volume 202 optimized for an intermediate pressure
and temperature at the combustion chamber 122) during an
intermediate steady state or transient operating condition of the
propulsion system 102. The outlet walls 220 articulate to
substantially block the first array 166 and the third array 170 to
prevent combustion gases 138 from the second array 168 from
backflowing into the first array 166 and third array 170 of
combustors 112. The outer walls 220 may further substantially block
a flow of oxidizer 195 through the first array 166 and the third
array 170 from flowing downstream to the exhaust section 106 while
enabling a flow of combustion gases 138 from the second array
168.
[0092] Although each array is described and shown in a certain
order relative to other arrays from the longitudinal centerline
116, it should be appreciated that the arrays 166, 168, 170 may be
arranged in other orders relative to the volumes of the combustion
chamber 122 defined. Still further, although certain operating
conditions are described in the context of aircraft landing/takeoff
cycles, it should be appreciated that the operating conditions may
include cycles specific to land- or marine-based power generation
gas turbine engines, auxiliary power units, turboprop or turboshaft
apparatuses, rockets, missiles, etc.
[0093] This written description uses examples to disclose the
invention, including the best mode, and also to enable any person
skilled in the art to practice the invention, including making and
using any devices or systems and performing any incorporated
methods. The patentable scope of the invention is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they include structural elements that do not
differ from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal languages of the claims.
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